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112-587: The engine produced a specific impulse ( I sp ) of 421 seconds (4.13 km/s) in a vacuum (or 200 seconds (2.0 km/s) at sea level) and had a mass of approximately 1,788 kilograms (3,942 lb). Five J-2 engines were used on the Saturn V's S-II second stage, and one J-2 was used on the S-IVB upper stage used on both the Saturn IB and Saturn V. Proposals also existed to use various numbers of J-2 engines in

224-409: A capacity of 118,931 cm (7,257.6 cu in). Both tanks were filled from a ground source prior to launch and the gaseous hydrogen tank was refilled during engine operation from the thrust chamber fuel inlet manifold for subsequent restart in third stage application. The control system included a pneumatic system and a solid-state electrical sequence controller packaged with spark exciters for

336-461: A device is suitable for its application, and will withstand the environmental conditions in which it is used. Normal operating temperature ranges are affected by several factors, such as the power dissipation of the device. These factors are used to define a "threshold temperature" of a device, i.e. its maximum normal operating temperature, and a maximum operating temperature beyond which the device will no longer function. Between these two temperatures,

448-417: A direct measure of the engine's effectiveness in converting propellant mass into forward momentum. The specific impulse in terms of propellant mass spent has units of distance per time, which is a notional velocity called the effective exhaust velocity . This is higher than the actual exhaust velocity because the mass of the combustion air is not being accounted for. Actual and effective exhaust velocity are

560-434: A given propellant, when paired with a given engine, can accelerate its own initial mass at 1 g. The longer it can accelerate its own mass, the more delta-V it delivers to the whole system. In other words, given a particular engine and a mass of a particular propellant, specific impulse measures for how long a time that engine can exert a continuous force (thrust) until fully burning that mass of propellant. A given mass of

672-526: A heavier engine with a higher specific impulse may not be as effective in gaining altitude, distance, or velocity as a lighter engine with a lower specific impulse, especially if the latter engine possesses a higher thrust-to-weight ratio . This is a significant reason for most rocket designs having multiple stages. The first stage is optimised for high thrust to boost the later stages with higher specific impulse into higher altitudes where they can perform more efficiently. The most common unit for specific impulse

784-411: A high velocity to the expelled combustion gases to produce thrust. The thrust chamber injector received the propellants under pressure from the turbopumps, then mixed them in a manner that produced the most efficient combustion. 614 hollow oxidizer posts were machined to form an integral part of the injector, with fuel nozzles (each swaged to the face of the injector) threaded through and installed over

896-514: A level of accumulated operational time almost eight times greater than the flight requirements. As successful single-engine tests moved toward their completion, integration tests of the propulsion system with the S-IVB accelerated with the availability of more production engines. The first operational flight, AS-201 , was scheduled in early 1966 for the Saturn IB using the S-IB first stage and the S-IVB as

1008-788: A material is "highly dependent on operating temperature", and creep analysis is thus an important part of design validation. Some of the effects of creep and thermal fatigue may be mitigated by integrating cooling systems into the device's design, reducing the peak temperature experienced by the metal. Commercial and retail products are manufactured to less stringent requirements than those for military and aerospace applications. For example, microprocessors produced by Intel Corporation are manufactured to three grades: commercial, industrial and extended. Because some devices generate heat during operation, they may require thermal management to ensure they are within their specified operating temperature range; specifically, that they are operating at or below

1120-737: A maximum of 6 hours, depending upon the number of Earth orbits required to attain the lunar window for translunar trajectory. Inspiration for the J-2 dates back to various NASA studies conducted in the late 1950s, of LH2-fuelled engines producing thrust of up to 665 kN (149,000 lb f ) following the success of the 67 kN (15,000 lb f ) RL-10 used on the Atlas-Centaur 's Centaur upper stage. As ever-heavier launch vehicles entered consideration, NASA began to look at engines producing thrusts of up to 890 kN (200,000 lb f ), with development being officially authorized following

1232-426: A maximum thrust of 5.7 N (1.3 lbf). Operating temperature An operating temperature is the allowable temperature range of the local ambient environment at which an electrical or mechanical device operates. The device will operate effectively within a specified temperature range which varies based on the device function and application context, and ranges from the minimum operating temperature to

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1344-417: A more energy-dense propellant can burn for a longer duration than some less energy-dense propellant made to exert the same force while burning in an engine. Different engine designs burning the same propellant may not be equally efficient at directing their propellant's energy into effective thrust. For all vehicles, specific impulse (impulse per unit weight-on-Earth of propellant) in seconds can be defined by

1456-417: A much higher specific impulse than rocket engines. For air-breathing engines, only the fuel mass is counted, not the mass of air passing through the engine. Air resistance and the engine's inability to keep a high specific impulse at a fast burn rate are limiting factors to the propellant consumption rate. If it were not for air resistance and the reduction of propellant during flight, specific impulse would be

1568-421: A much larger specific impulse than a rocket; for example a turbofan jet engine may have a specific impulse of 6,000 seconds or more at sea level whereas a rocket would be between 200 and 400 seconds. An air-breathing engine is thus much more propellant efficient than a rocket engine, because the air serves as reaction mass and oxidizer for combustion which does not have to be carried as propellant, and

1680-478: A new "Idle Mode" that produced little thrust for on-orbit maneuvering or to settle the fuel tanks on-orbit prior to a burn. During the experimental program, Rocketdyne also produced a small run of six pre-production models for testing, the J-2S . These were test fired many times between 1965 and 1972, for a total of 30,858 seconds burn time. In 1972 it became clear no follow-on orders for Saturn boosters were coming, and

1792-412: A number of valves to control the operation of the engine by changing the flow of propellant through the engine's components: The gas generator system consisted of the gas generator, gas generator control valve, turbine exhaust system and exhaust manifold, heat exchanger, and oxidizer turbine bypass valve. The gas generator itself was welded to the fuel pump turbine manifold, making it an integral part of

1904-400: A rocket can be defined in terms of thrust per unit mass flow of propellant. This is an equally valid (and in some ways somewhat simpler) way of defining the effectiveness of a rocket propellant. For a rocket, the specific impulse defined in this way is simply the effective exhaust velocity relative to the rocket, v e . "In actual rocket nozzles, the exhaust velocity is not really uniform over

2016-407: A system fan)", though in "a properly designed system, this feature should never become active". Cooling and other thermal management techniques may affect performance and noise level. Noise mitigation strategies may be required in residential applications to ensure that the noise level does not become uncomfortable. Battery service life and efficacy is affected by operating temperature. Efficacy

2128-484: Is a measure of how efficiently a reaction mass engine, such as a rocket using propellant or a jet engine using fuel, generates thrust . A propulsion system with a higher specific impulse uses the mass of the propellant more efficiently. In the case of a rocket, this means less propellant needed for a given delta- v , so that the vehicle attached to the engine can more efficiently gain altitude and velocity. For engines like cold gas thrusters whose reaction mass

2240-490: Is also ionized, which would interfere with radio communication with the rocket. Nuclear thermal rocket engines differ from conventional rocket engines in that energy is supplied to the propellants by an external nuclear heat source instead of the heat of combustion . The nuclear rocket typically operates by passing liquid hydrogen gas through an operating nuclear reactor. Testing in the 1960s yielded specific impulses of about 850 seconds (8,340 m/s), about twice that of

2352-450: Is also valid for air-breathing jet engines, but is rarely used in practice. (Note that different symbols are sometimes used; for example, c is also sometimes seen for exhaust velocity. While the symbol I sp {\displaystyle I_{\text{sp}}} might logically be used for specific impulse in units of (N·s )/(m·kg); to avoid confusion, it is desirable to reserve this for specific impulse measured in seconds.) It

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2464-460: Is determined by comparing the service life achieved by the battery as a percentage of its service life achieved at 20 °C (68 °F) versus temperature. Ohmic load and operating temperature often jointly determine a battery's discharge rate. Moreover, if the expected operating temperature for a primary battery deviates from the typical 10 °C to 25 °C (50 to 77 °F) range, then operating temperature "will often have an influence on

2576-434: Is impractical. Lithium and fluorine are both extremely corrosive, lithium ignites on contact with air, fluorine ignites on contact with most fuels, and hydrogen, while not hypergolic, is an explosive hazard. Fluorine and the hydrogen fluoride (HF) in the exhaust are very toxic, which damages the environment, makes work around the launch pad difficult, and makes getting a launch license that much more difficult. The rocket exhaust

2688-497: Is inversely proportional to specific fuel consumption (SFC) by the relationship I sp = 1/( g o ·SFC) for SFC in kg/(N·s) and I sp = 3600/SFC for SFC in lb/(lbf·hr). An example of a specific impulse measured in time is 453 seconds, which is equivalent to an effective exhaust velocity of 4.440 km/s (14,570 ft/s), for the RS-25 engines when operating in a vacuum. An air-breathing jet engine typically has

2800-423: Is needed to produce a given thrust for a given time and the more efficient the propellant is. This should not be confused with the physics concept of energy efficiency , which can decrease as specific impulse increases, since propulsion systems that give high specific impulse require high energy to do so. Thrust and specific impulse should not be confused. Thrust is the force supplied by the engine and depends on

2912-421: Is only the fuel they carry, specific impulse is exactly proportional to the effective exhaust gas velocity. In an atmospheric context, specific impulse can include the contribution to impulse provided by the mass of external air that is accelerated by the engine, such as by fuel combustion or by external propeller. Jet engines and turbofans breathe external air for both combustion and bypass, and therefore have

3024-408: Is proportional to the effective exhaust velocity. A spacecraft without propulsion follows an orbit determined by its trajectory and any gravitational field. Deviations from the corresponding velocity pattern (these are called Δ v ) are achieved by sending exhaust mass in the direction opposite to that of the desired velocity change. When an engine is run within the atmosphere, the exhaust velocity

3136-410: Is reduced by atmospheric pressure, in turn reducing specific impulse. This is a reduction in the effective exhaust velocity, versus the actual exhaust velocity achieved in vacuum conditions. In the case of gas-generator cycle rocket engines, more than one exhaust gas stream is present as turbopump exhaust gas exits through a separate nozzle. Calculating the effective exhaust velocity requires averaging

3248-409: Is related to the thrust , or forward force on the rocket by the equation: F thrust = v e ⋅ m ˙ , {\displaystyle F_{\text{thrust}}=v_{\text{e}}\cdot {\dot {m}},} where m ˙ {\displaystyle {\dot {m}}} is the propellant mass flow rate, which is the rate of decrease of

3360-455: Is that it may be used for rockets, where all the reaction mass is carried on board, as well as airplanes, where most of the reaction mass is taken from the atmosphere. In addition, giving the result as a unit of time makes the result easily comparable between calculations in SI units, imperial units, US customary units or other unit framework. The English unit pound mass is more commonly used than

3472-401: Is the product of the average specific gravity of a given propellant mixture and the specific impulse. While less important than the specific impulse, it is an important measure in launch vehicle design, as a low specific impulse implies that bigger tanks will be required to store the propellant, which in turn will have a detrimental effect on the launch vehicle's mass ratio . Specific impulse

Rocketdyne J-2 - Misplaced Pages Continue

3584-481: Is the second, as values are identical regardless of whether the calculations are done in SI , imperial , or US customary units. Nearly all manufacturers quote their engine performance in seconds, and the unit is also useful for specifying aircraft engine performance. The use of metres per second to specify effective exhaust velocity is also reasonably common. The unit is intuitive when describing rocket engines, although

3696-414: Is used, impulse is divided by propellant weight (weight is a measure of force), resulting in units of time (seconds). These two formulations differ from each other by the standard gravitational acceleration ( g 0 ) at the surface of the earth. The rate of change of momentum of a rocket (including its propellant) per unit time is equal to the thrust. The higher the specific impulse, the less propellant

3808-626: The J-2T-200k that provided 890 kN (200,000 lbf) thrust, allowing it to be "dropped in" to the existing S-II and S-IVB stages, and the J-2T-250k of 1,100 kN (250,000 lbf). Like the J-2S, work on the J-2T had progressed to a lengthy series of ground-based test runs, but further development ended in the post-Apollo draw-down. What became a different engine with a similar name, called

3920-681: The J-2X , was chosen in 2007 for the Project Constellation crewed lunar landing program. A single J-2X engine, generating 1,310 kN (294,000 lbf) of thrust, was to be used to power the Earth Departure Stage (EDS). NASA began construction of a new test stand for altitude testing of J-2X engines at Stennis Space Center (SSC) on 23 August 2007. Between December 2007 and May 2008, nine tests of heritage J-2 engine components were conducted at SSC in preparation for

4032-815: The United States Department of Defense has defined the United States Military Standard for all products used by the United States Armed Forces. A product's environmental design and test limits to the conditions that it will undergo throughout its service life are specified in MIL-STD-810 , the Department of Defense Test Method Standard for Environmental Engineering Considerations and Laboratory Tests . The MIL-STD-810G standard specifies that

4144-416: The basal body temperature , is achieved during sleep. In women, it is affected by ovulation, causing a biphasic pattern which may be used as a component of fertility awareness . In humans, the hypothalamus regulates metabolism , and hence the basal metabolic rate . Amongst its functions is the regulation of body temperature. The core body temperature is also one of the classic phase markers for measuring

4256-404: The effective exhaust velocity while reducing the actual exhaust velocity. Again, this is because the mass of the air is not counted in the specific impulse calculation, thus attributing all of the thrust momentum to the mass of the fuel component of the exhaust, and omitting the reaction mass, inert gas, and effect of driven fans on overall engine efficiency from consideration. Essentially,

4368-675: The maximum operating temperature (or peak operating temperature ). Outside this range of safe operating temperatures the device may fail. It is one component of reliability engineering . Similarly, biological systems have a viable temperature range, which might be referred to as an "operating temperature". Most semiconductor devices are manufactured in several temperature grades. Broadly accepted grades are: Nevertheless, each manufacturer defines its own temperature grades so designers must pay attention to datasheet specifications. For example, Maxim Integrated uses five temperature grades for its products: The use of such grades ensures that

4480-487: The "operating temperature stabilization is attained when the temperature of the functioning part(s) of the test item considered to have the longest thermal lag is changing at a rate of no more than 2.0 °C (3.6 °F) per hour." It also specifies procedures to assess the performance of materials to extreme temperature loads . Military engine turbine blades experience two significant deformation stresses during normal service, creep and thermal fatigue . Creep life of

4592-624: The 1959 report of the Saturn Vehicle Evaluation Committee . A source evaluation board was formed to nominate a contractor from five bidding companies, and approval was given on 1 June 1960 for Rocketdyne to begin development of a "high-energy rocket engine, fuelled by LOX and hydrogen, to be known as the J-2". The final contract, awarded in September 1960, was the first to explicitly require the design "insure maximum safety for crewed flight ." Rocketdyne launched

Rocketdyne J-2 - Misplaced Pages Continue

4704-449: The ASI chamber that passed through the center of the thrust chamber injector. After a delay of 1, 3, or 8 seconds, during which time fuel was circulated through the thrust chamber to condition the engine for start, the start tank discharge valve was opened to initiate turbine spin. The length of the fuel lead was dependent upon the length of the Saturn V first stage boost phase. When the engine

4816-422: The ASI. The ASI operated continuously during entire engine firing, was uncooled, and was capable of multiple reignitions under all environmental conditions. Thrust was transmitted through the gimbal (mounted to the injector and oxidizer dome assembly and the vehicle's thrust structure), which consisted of a compact, highly loaded (140,000 kPa) universal joint consisting of a spherical, socket-type bearing. This

4928-523: The Saturn V S-IVB third stage. The first burn, lasting about two minutes, placed the Apollo spacecraft into a low Earth parking orbit . After the crew verified that the spacecraft was operating nominally, the J-2 was re-ignited for translunar injection , a 6.5 minute burn which accelerated the vehicle to a course for the Moon . The J-2's thrust chamber assembly served as a mount for all engine components, and

5040-693: The Space Shuttle engines. A variety of other rocket propulsion methods, such as ion thrusters , give much higher specific impulse but with much lower thrust; for example the Hall-effect thruster on the SMART-1 satellite has a specific impulse of 1,640 s (16.1 km/s) but a maximum thrust of only 68 mN (0.015 lbf). The variable specific impulse magnetoplasma rocket (VASIMR) engine currently in development will theoretically yield 20 to 300 km/s (66,000 to 984,000 ft/s), and

5152-408: The actual exhaust speed is much lower, so the kinetic energy the exhaust carries away is lower and thus the jet engine uses far less energy to generate thrust. While the actual exhaust velocity is lower for air-breathing engines, the effective exhaust velocity is very high for jet engines. This is because the effective exhaust velocity calculation assumes that the carried propellant is providing all

5264-479: The ambient temperature, and for integrated circuits , is given by the equation: in which T J is the junction temperature in °C, T a is the ambient temperature in °C, P D is the power dissipation of the integrated circuit in W , and R ja is the junction to ambient thermal resistance in °C/W. Electrical and mechanical devices used in military and aerospace applications may need to endure greater environmental variability, including temperature range. In

5376-419: The amount of reaction mass flowing through the engine. Specific impulse measures the impulse produced per unit of propellant and is proportional to the exhaust velocity. Thrust and specific impulse are related by the design and propellants of the engine in question, but this relationship is tenuous. For example, LH 2 /LO 2 bipropellant produces higher I sp but lower thrust than RP-1 / LO 2 due to

5488-427: The burned fuel. Next, inert gases in the atmosphere absorb heat from combustion, and through the resulting expansion provide additional thrust. Lastly, for turbofans and other designs there is even more thrust created by pushing against intake air which never sees combustion directly. These all combine to allow a better match between the airspeed and the exhaust speed, which saves energy/propellant and enormously increases

5600-399: The combustion chamber approximately halfway between the throat and the nozzle exit. Exhaust gases passed through the heat exchanger and exhaust into the main combustion chamber through 180 triangular openings between the tubes of the combustion chamber. The heat exchanger was a shell assembly, consisting of a duct, bellows, flanges, and coils. It was mounted in the turbine exhaust duct between

5712-424: The combustion chamber instead of a separate burner. In addition to removing parts from the engine, it also reduced the difficulty of starting up the engine and properly timing various combustors. Additional changes included a throttling system for wider mission flexibility, which also required a variable mixture system to properly mix the fuel and oxygen for a variety of different operating pressures. It also included

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5824-418: The control valve to the injector assembly and into the combustor outlet, before being directed to the fuel turbine and then to the oxidizer turbine. The turbine exhaust ducting and turbine exhaust hoods were of welded sheet metal construction. Flanges utilizing dual seals were used at component connections. The exhaust ducting conducted turbine exhaust gases to the thrust chamber exhaust manifold which encircled

5936-474: The definition of specific impulse as impulse per unit mass of propellant. Specific fuel consumption is inversely proportional to specific impulse and has units of g/(kN·s) or lb/(lbf·h). Specific fuel consumption is used extensively for describing the performance of air-breathing jet engines. Specific impulse, measured in seconds, can be thought of as how many seconds one kilogram of fuel can produce one kilogram of thrust. Or, more precisely, how many seconds

6048-553: The design of the J-2X engine. The new J-2X is designed to be more efficient and simpler to build than its Apollo J-2 predecessor, and cost less than the Space Shuttle Main Engine (SSME). Design differences include the removal of beryllium , modern electronics, a centrifugal turbo pump versus the axial turbo pump of the J-2, a different chamber and nozzle expansion ratios, a channel-walled combustion chamber versus

6160-416: The development of the J-2 with an analytical computer model that simulated engine operations and aided in establishing design configurations. The model was supported by a full-sized mockup which was used throughout development to judge the positioning of the engine's components. The first experimental component, the engine's injector , was produced within two months of the contract being awarded, and testing of

6272-641: The development process, assisting in the design of the engine's electrical and mechanical systems. Contracts were signed between NASA and Rocketdyne in the summer of 1962, requiring 55 J-2 engines to be produced to support the final designs for the Saturn rockets , which required five engines for each S-II second stage of the Saturn V and one engine for each S-IVB Saturn IB and Saturn V third stage. The J-2 entered production in May 1963, with concurrent testing programs continuing to run at Rocketdyne and at MSFC during

6384-411: The device will operate at a non-peak level. For instance, a resistor may have a threshold temperature of 70 °C (158 °F) and a maximum temperature of 155 °C (311 °F), between which it exhibits a thermal derating . For electrical devices, the operating temperature may be the junction temperature (T J ) of the semiconductor in the device. The junction temperature is affected by

6496-492: The effective exhaust speed of the engines may be significantly different from the actual exhaust speed, especially in gas-generator cycle engines. For airbreathing jet engines , the effective exhaust velocity is not physically meaningful, although it can be used for comparison purposes. Metres per second are numerically equivalent to newton-seconds per kg (N·s/kg), and SI measurements of specific impulse can be written in terms of either units interchangeably. This unit highlights

6608-424: The engine components, with the capability of transmitting signals to a ground recording system or a telemetry system, or both. The instrumentation system was designed for use throughout the life of the engine, from the first static acceptance firing to its ultimate vehicle flight. The auxiliary package was designed for use during early vehicle flights. It may be deleted from the basic engine instrumentation system after

6720-588: The engine's components began at Rocketdyne's Santa Susana Field Laboratory in November 1960. Other test facilities, including a vacuum chamber and full-size engine test stand, were used during the development, with the engine's turbopumps entering testing in November 1961, the ignition system in early 1962, and the first prototype engine running a complete 250-second test run in October 1962. In addition to flight hardware, five engine simulators were also used during

6832-400: The engine, with two uprated versions being used by NASA in the later flights of the Apollo program. An experimental program to improve the performance of the J-2 started in 1964 as the J-2X (not to be confused with a later variant by the same name ). The main change to the original J-2 design was a change from the gas generator cycle to a tap-off cycle that supplied hot gas from a tap on

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6944-638: The entire exit cross section and such velocity profiles are difficult to measure accurately. A uniform axial velocity, v e , is assumed for all calculations which employ one-dimensional problem descriptions. This effective exhaust velocity represents an average or mass equivalent velocity at which propellant is being ejected from the rocket vehicle." The two definitions of specific impulse are proportional to one another, and related to each other by: v e = g 0 ⋅ I sp , {\displaystyle v_{\text{e}}=g_{0}\cdot I_{\text{sp}},} where This equation

7056-435: The exhaust gases having a lower density and higher velocity ( H 2 O vs CO 2 and H 2 O). In many cases, propulsion systems with very high specific impulse—some ion thrusters reach 10,000 seconds—produce low thrust. When calculating specific impulse, only propellant carried with the vehicle before use is counted. For a chemical rocket, the propellant mass therefore would include both fuel and oxidizer . In rocketry,

7168-455: The following equation: F thrust = g 0 ⋅ I sp ⋅ m ˙ , {\displaystyle F_{\text{thrust}}=g_{0}\cdot I_{\text{sp}}\cdot {\dot {m}},} where: I sp in seconds is the amount of time a rocket engine can generate thrust, given a quantity of propellant whose weight is equal to the engine's thrust. The advantage of this formulation

7280-507: The fuel and oxidizer high-pressure ducts. The flowmeters measured propellant flowrates in the high-pressure propellant ducts. The four-vane rotor in the hydrogen system produced four electrical impulses per revolution and turned approximately 3,700 rpm at nominal flow. The six-vane rotor in the LOX system produced six electrical impulses per revolution and turned at approximately 2,600 rpm at nominal flow. The propellant feed system required

7392-416: The fuel inlet manifold. Power from the turbine was transmitted by means of a one-piece shaft to the pump. The velocity of the LOX was increased through the inducer and impeller. As the LOX entered the outlet volute, velocity was converted to pressure and the LOX was discharged into the outlet duct at high pressure. The fuel and oxidizer flowmeters were helical-vaned, rotor-type flowmeters. They were located in

7504-419: The fuel orifices which were concentric with the oxidizer orifices. The propellants were injected uniformly to ensure satisfactory combustion. The injector and oxidizer dome assembly was located at the top of the thrust chamber. The dome provided a manifold for the distribution of the LOX to the injector and served as a mount for the gimbal bearing and the augmented spark igniter. The augmented spark igniter (ASI)

7616-422: The fuel turbopump assembly. It produced hot gases to drive the fuel and oxidizer turbines and consisted of a combustor containing two spark plugs, a control valve containing fuel and oxidizer ports, and an injector assembly. When engine start was initiated, the spark exciters in the electrical control package were energized, providing energy to the spark plugs in the gas generator combustor. Propellants flowed through

7728-446: The fuel. Fuel entered from a manifold , located midway between the thrust chamber throat and the exit, at a pressure of more than 6,900 kPa (1,000 psi). In cooling the chamber, the fuel made a one-half pass downward through 180 tubes and was returned in a full pass up to the thrust chamber injector through 360 tubes. Once propellants passed through the injector, they were ignited by the augmented spark igniter and burned to impart

7840-441: The gas generator and the thrust chamber spark plugs, plus interconnecting electrical cabling and pneumatic lines, in addition to the flight instrumentation system. The pneumatic system consisted of a high-pressure helium gas storage tank, a regulator to reduce the pressure to a usable level, and electrical solenoid control valves to direct the central gas to the various pneumatically controlled valves. The electrical sequence controller

7952-459: The gas generator and two in the augmented spark igniter for ignition of the propellants. Next, two solenoid valves were actuated; one for helium control, and one for ignition phase control. Helium was routed to hold the propellant bleed valves closed and to purge the thrust chamber LOX dome, the LOX pump intermediate seal, and the gas generator oxidizer passage. In addition, the main fuel and ASI oxidizer valves were opened, creating an ignition flame in

8064-404: The gaseous helium tank was not required because the original ground-fill supply was sufficient for three starts). Prior to engine restart, the stage ullage rockets were fired to settle the propellants in the stage propellant tanks, ensuring a liquid head to the turbopump inlets. In addition, the engine propellant bleed valves were opened, the stage recirculation valve was opened, the stage prevalve

8176-399: The initial J-2X gas generator design. The completion of a second round of successful gas generator tests was announced on September 21, 2010. Project Constellation was cancelled by President Barack Obama on October 11, 2010, but development of the J-2X has continued for its potential as the second stage engine for the new, heavy-lift Space Launch System . The first hot-fire test of the J-2X

8288-504: The initial phase of engine operation, the gaseous hydrogen start tank would be recharged in those engines having a restart requirement. The hydrogen tank was repressurized by tapping off a controlled mixture of LH 2 from the thrust chamber fuel inlet manifold and warmer hydrogen from the thrust chamber fuel injection manifold just before entering the injector. During mainstage operation, engine thrust could be varied between 780 and 1,000 kilonewtons (175,000 and 225,000 lbf) by actuating

8400-430: The main fuel, main oxidizer, gas generator control, and augmented spark igniter valves. The oxidizer turbine bypass valve and propellant bleed valves opened and the gas generator and LOX dome purges were initiated. To provide third stage restart capability for the Saturn V, the J-2 gaseous hydrogen start tank was refilled in 60 seconds during the previous firing after the engine had reached steady-state operation (refill of

8512-650: The manufacturing run. The first production engine, delivered in April 1964, went for static tests on the S-IVB test stage at the Douglas test facility near Sacramento, California and underwent its first full-duration (410 seconds) static test in December 1964. Testing continued until January 1966, with one engine in particular igniting successfully in 30 successive firings, including five tests at full duration of 470 seconds each. The total firing time of 3774 seconds represented

8624-435: The maximum operating temperature of the device. Cooling a microprocessor mounted in a typical commercial or retail configuration requires "a heatsink properly mounted to the processor, and effective airflow through the system chassis". Systems are designed to protect the processor from unusual operating conditions, such as "higher than normal ambient air temperatures or failure of a system thermal management component (such as

8736-490: The momentum of engine exhaust includes a lot more than just fuel, but specific impulse calculation ignores everything but the fuel. Even though the effective exhaust velocity for an air-breathing engine seems nonsensical in the context of actual exhaust velocity, this is still useful for comparing absolute fuel efficiency of different engines. A related measure, the density specific impulse , sometimes also referred to as Density Impulse and usually abbreviated as I s d

8848-465: The only reaction mass is the propellant, so the specific impulse is calculated using an alternative method, giving results with units of seconds. Specific impulse is defined as the thrust integrated over time per unit weight -on-Earth of the propellant: I sp = v e g 0 , {\displaystyle I_{\text{sp}}={\frac {v_{\text{e}}}{g_{0}}},} where In rockets, due to atmospheric effects,

8960-417: The oxidizer posts in concentric rings. The injector face was porous, being formed from layers of stainless steel wire mesh, and was welded at its periphery to the injector body. The injector received LOX through the dome manifold and injected it through the oxidizer posts into the combustion area of the thrust chamber, while fuel was received from the upper fuel manifold in the thrust chamber and injected through

9072-417: The oxidizer turbine discharge manifold and the thrust chamber. It heated and expanded helium gas for use in the third stage or converted LOX to gaseous oxygen for the second stage for maintaining vehicle oxidizer tank pressurization. During engine operation, either LOX was tapped off the oxidizer high-pressure duct or helium was provided from the vehicle stage and routed to the heat exchanger coils. This system

9184-403: The pressure of the LOX and pumps it through high-pressure ducts to the thrust chamber. The pump operated at 8,600 rpm at a discharge pressure of 7,400 kPa (1,080 psi) (absolute) and developed 1,600 kW (2,200 bhp). The pump and its two turbine wheels are mounted on a common shaft. Power for operating the oxidizer turbopump was provided by a high-speed, two-stage turbine which

9296-559: The program shut down. NASA did consider using the J-2S on a number of different missions, including powering the Space Shuttle in a number of early designs as well as on the Comet HLLV . While work on the J-2S continued, NASA also funded a design effort to use the J-2S turbomachinery and plumbing to a toroidal combustion chamber with a new aerospike nozzle. This would improve performance even further. Two versions were built,

9408-438: The propellant utilization valve to increase or decrease oxidizer flow. This was beneficial to flight trajectories and for overall mission performance to make greater payloads possible. When the engine cutoff signal was received by the electrical control package, it de-energized the main-stage and ignition phase solenoid valves and energized the helium control solenoid de-energizer timer. This, in turn, permitted closing pressure to

9520-445: The propulsion system has established its reliability during research and development vehicle flights. It contains sufficient flexibility to provide for deletion, substitution, or addition of parameters deemed necessary as a result of additional testing. Eventual deletion of the auxiliary package will not interfere with the measurement capability of the primary package. Start sequence was initiated by supplying energy to two spark plugs in

9632-414: The reaction mass and all the thrust. Hence effective exhaust velocity is not physically meaningful for air-breathing engines; nevertheless, it is useful for comparison with other types of engines. The highest specific impulse for a chemical propellant ever test-fired in a rocket engine was 542 seconds (5.32 km/s) with a tripropellant of lithium , fluorine , and hydrogen . However, this combination

9744-422: The same in rocket engines operating in a vacuum. The amount of propellant can be measured either in units of mass or weight. If mass is used, specific impulse is an impulse per unit of mass, which dimensional analysis shows to have units of speed, specifically the effective exhaust velocity . As the SI system is mass-based, this type of analysis is usually done in meters per second. If a force-based unit system

9856-533: The second stage. The first all-up test of a complete S-IVB, including its single J-2, in July 1965 was inconclusive when a component malfunction in one of the pneumatic consoles prematurely ended the test after a successful propellant loading and automatic countdown. Confidence in the design was regained in August, however, when the same stage, S-IVB-201, performed flawlessly on a full-duration firing of 452 seconds, which

9968-574: The slug, and when using pounds per second for mass flow rate, it is more convenient to express standard gravity as 1 pound-force per pound-mass. Note that this is equivalent to 32.17405 ft/s2, but expressed in more convenient units. This gives: F thrust = I sp ⋅ m ˙ ⋅ ( 1 l b f l b m ) . {\displaystyle F_{\text{thrust}}=I_{\text{sp}}\cdot {\dot {m}}\cdot \left(1\mathrm {\frac {lbf}{lbm}} \right).} In rocketry,

10080-490: The specific impulse varies with altitude, reaching a maximum in a vacuum. This is because the exhaust velocity isn't simply a function of the chamber pressure, but is a function of the difference between the interior and exterior of the combustion chamber . Values are usually given for operation at sea level ("sl") or in a vacuum ("vac"). Because of the geocentric factor of g 0 in the equation for specific impulse, many prefer an alternative definition. The specific impulse of

10192-406: The system automatically reset for a subsequent restart. The flight instrumentation system was composed of a primary instrumentation package and an auxiliary package. The primary package instrumentation measures those parameters critical to all engine static firings and subsequent vehicle launches. These include some 70 parameters such as pressures, temperatures, flows, speeds, and valve positions for

10304-409: The thrust chamber, was a turbine-driven, axial flow pumping unit consisting of an inducer, a seven-stage rotor, and a stator assembly. It was a high-speed pump operating at 27,000 rpm, and was designed to increase hydrogen pressure from 210 to 8,450 kPa (30 to 1,225 psi) (absolute) through high-pressure ducting at a flowrate which develops 5,800 kW (7,800 bhp). Power for operating

10416-401: The timing of an individual's Circadian rhythm . Changes to the normal human body temperature may result in discomfort. The most common such change is a fever , a temporary elevation of the body's thermoregulatory set-point, typically by about 1–2 °C (1.8–3.6 °F). Hyperthermia is an acute condition caused by the body absorbing more heat than it can dissipate, whereas hypothermia

10528-543: The tube-welded chamber of the J-2, a redesign of all the electronics, supersonic injection and the use of 21st-century joining techniques. On July 16, 2007 NASA officially announced the award to Pratt & Whitney Rocketdyne , Inc. of a $ 1.2 billion contract "for design, development, testing and evaluation of the J-2X engine" intended to power the upper stages of the Ares I and Ares V launch vehicles. On Sept. 8, 2008 Pratt & Whitney Rocketdyne announced successful testing of

10640-411: The turbine through the exhaust ducting. Three dynamic seals in series prevented the pump fluid and turbine gas from mixing. Power from the turbine was transmitted to the pump by means of a one-piece shaft. The oxidizer turbopump was mounted on the thrust chamber diametrically opposite the fuel turbopump. It was a single-stage centrifugal pump with direct turbine drive . The oxidizer turbopump increases

10752-400: The turbopump operation, hot gas entered the nozzles and, in turn, the first stage turbine wheel. After passing through the first stage turbine wheel, the gas was redirected by the stator blades and entered the second stage turbine wheel. The gas then left the turbine through exhaust ducting, passed through the heat exchanger, and exhausted into the thrust chamber through a manifold directly above

10864-435: The turbopump was provided by a high-speed, two-stage turbine. Hot gas from the gas generator was routed to the turbine inlet manifold which distributed the gas to the inlet nozzles where it was expanded and directed at a high velocity into the first stage turbine wheel. After passing through the first stage turbine wheel, the gas was redirected through a ring of stator blades and enters the second stage turbine wheel. The gas left

10976-415: The two mass flows as well as accounting for any atmospheric pressure. For air-breathing jet engines, particularly turbofans , the actual exhaust velocity and the effective exhaust velocity are different by orders of magnitude. This happens for several reasons. First, a good deal of additional momentum is obtained by using air as reaction mass, such that combustion products in the exhaust have more mass than

11088-406: The type of battery selected for the application". Energy reclamation from partially depleted lithium sulfur dioxide battery has been shown to improve when "appropriately increasing the battery operating temperature". Mammals attempt to maintain a comfortable body temperature under various conditions by thermoregulation , part of mammalian homeostasis . The lowest normal temperature of a mammal,

11200-525: The upper stages of an even larger rocket, the planned Nova . The J-2 was America's largest production LH2-fuelled rocket engine before the RS-25 . A modernized version of the engine, the J-2X , was considered for use on the Earth Departure Stage of NASA's Space Shuttle replacement, the Space Launch System . Unlike most liquid-fueled rocket engines in service at the time, the J-2 was designed to be restarted once after shutdown when flown on

11312-446: The vehicle's mass. A rocket must carry all its propellant with it, so the mass of the unburned propellant must be accelerated along with the rocket itself. Minimizing the mass of propellant required to achieve a given change in velocity is crucial to building effective rockets. The Tsiolkovsky rocket equation shows that for a rocket with a given empty mass and a given amount of propellant, the total change in velocity it can accomplish

11424-490: The vehicle. The propellant feed system consists of separate fuel and oxidizer turbopumps (the bearings of which were lubricated by the fluid being pumped because the extremely low operating temperature of the engine precluded use of lubricants or other fluids), several valves (including the main fuel valve, main oxidizer valve, propellant utilization valve and fuel and oxidizer bleed valves), fuel and oxidizer flowmeters, and interconnecting lines. The fuel turbopump, mounted on

11536-447: Was a completely self-contained, solid-state system, requiring only DC power and start and stop command signals. Pre-start status of all critical engine control functions was monitored in order to provide an "engine ready" signal. Upon obtaining "engine ready" and "start" signals, solenoid control valves were energized in a precisely timed sequence to bring the engine through ignition, transition, and into main-stage operation. After shutdown,

11648-411: Was closed, and a LOX and LH 2 circulation was effected through the engine bleed system for five minutes to condition the engine to the proper temperature to ensure proper engine operation. Engine restart was initiated after the "engine ready" signal was received from the stage. This was similar to the initial "engine ready". The hold time between cutoff and restart was from a minimum of 1.5 hours to

11760-426: Was composed of the thrust chamber body, injector and dome assembly, gimbal bearing assembly, and augmented spark igniter. The thrust chamber was constructed of 0.30 millimetres (0.012 in) thick stainless steel tubes, stacked longitudinally and furnace-brazed to form a single unit. The chamber was bell-shaped with a 27.5:1 expansion area ratio for efficient operation at altitude, and was regeneratively cooled by

11872-425: Was covered with a Teflon/fiberglass coating that provided a dry, low-friction bearing surface. The gimbal included a lateral adjustment device for aligning the combustion chamber with the vehicle, so that, in addition to transmitting the thrust from the injector assembly to the vehicle thrust structure, the gimbal also provided a pivot bearing for deflection of the thrust vector, thus providing flight attitude control of

11984-403: Was driven by the exhaust gases from the gas generator. The turbines of the oxidizer and fuel turbopumps were connected in a series by exhaust ducting that directed the discharged exhaust gas from the fuel turbopump turbine to the inlet of the oxidizer turbopump turbine manifold. One static and two dynamic seals in series prevented the turbopump oxidizer fluid and turbine gas from mixing. Beginning

12096-522: Was made up of an integral helium and hydrogen start tank, which contained the hydrogen and helium gases for starting and operating the engine. The gaseous hydrogen imparted initial spin to the turbines and pumps prior to gas generator combustion, and the helium was used in the control system to sequence the engine valves. The spherical helium tank was positioned inside the hydrogen tank to minimize engine complexity. It held 16,000 cm (1,000 cu in) of helium. The larger spherical hydrogen gas tank had

12208-506: Was mounted to the injector face and provided the flame to ignite the propellants in the combustion chamber. When engine start was initiated, the spark exciters energized two spark plugs mounted in the side of the combustion chamber. Simultaneously, the control system started the initial flow of oxidizer and fuel to the spark igniter. As the oxidizer and fuel entered the combustion chamber of the ASI, they mixed and were ignited, with proper ignition being monitored by an ignition monitor mounted in

12320-579: Was scheduled for late June, 2011. On November 9, 2011 NASA conducted a successful firing of the J-2X engine of 499.97 seconds in duration. On February 27, 2013 NASA continued testing of the J-2X engine of 550 seconds in duration at NASA's Stennis Space Center. [REDACTED]  This article incorporates public domain material from websites or documents of the National Aeronautics and Space Administration . Specific impulse Specific impulse (usually abbreviated I sp )

12432-564: Was the first engine test sequence to be controlled entirely by computers. The J-2 was cleared for flight and, on 26 February 1966, AS-201 went through a flawless launch. In July 1966, NASA confirmed J-2 production contracts through 1968, by which time Rocketdyne agreed to finish deliveries of 155 J-2 engines, with each engine undergoing a flight qualification firing at the Santa Susana Field Laboratory before delivery to NASA. Reliability and development testing continued on

12544-467: Was used in the S-II stage, a 1-second fuel lead was necessary. The S-IVB, on the other hand, utilized a 1-second fuel lead for its initial start and an 8-second fuel lead for its restart. After an interval of 0.450 seconds, the start tank discharge valve was closed and a mainstage control solenoid was actuated to: Energy in the spark plugs was cut off and the engine was operating at rated thrust. During

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